The present invention is directed to turbo-machines and, more particularly, to multistage axial or radial gas flow compressors and turbines and systems employing such turbo-machines.
It is known that the efficiency of turbo-machines, such as compressors and gas turbines, may be substantially improved by operation in a manner which approaches isothermal conditions. This essentially means that the temperature of the gas as it moves between successive stages of the turbo-machine is adjusted so that the inlet temperature of the gas at each successive stage is maintained at about the same temperature as at the inlet of the preceding stage. This is in contrast to adiabatic operation in which the temperature of the gas changes between the successive stages due to the compression or expansion of the gas as it moves through each successive stage of the turbo-machine.
Maintenance of a constant temperature at the inlet of each successive stage may be accomplished in several different ways. In a purely isothermal gas turbine, fuel injectors and temperature sensors may be positioned in each stage so that the correct amount of fuel is injected into and burned in each stage as is needed to ensure that the temperature of the gas in the gas turbine is re-elevated to substantially the temperature at which it entered that stage prior to discharge from the stage and introduction to the next succeeding stage. This is shown for example in U.S. Pat. No. 4,197,700 (Jahnig). In a purely isothermal compressor, a coolant may be introduced into each stage, for example through the stator blades of an axial compressor, to reduce the temperature of the gas to substantially the same temperature at which it was introduced to that stage to ensure that the temperature of the gas which is discharged from the stage and introduced to the next stage is at substantially the same temperature. Combustion chambers or intercoolers have also been employed between stages to add or remove heat and alter the gas temperature so that the gas entering each of the respective stages is at substantially the same temperature.
Substantial improvements in efficiency may also be achieved in particular in compressors through the use of relatively low temperature coolants, such as sea water which is taken from below the thermocline. Such sea water will typically be about 40.degree. F. which is sufficient to maintain a temperature of about 45.degree. F. to the intake of each stage of an isothermal compressor.
It would also be desirable to design, for example, the first stage of the turbo-machine to achieve the maximum efficiency from a design standpoint when the turbo-machine is in normal operation. Normal operation means that each stage would have a given shaft speed, pressure ratio, temperature ratio, gas density ratio, and the type of operation in each stage would be the same, e.g. isothermal, adiabatic, etc. This optimum efficiency stage could then act as a master stage which would serve as a model for the construction of each of the subsequent stages. In the present invention a formula has been discovered for the sizing of each subsequent stage once an optimum efficiency master stage has been designed which will maximize the optimum efficiency of each subsequent stage so that it has substantially the same optimum efficiency as the optimum efficiency master stage.
It has also been discovered that the sizing formula of the present invention is applicable to all turbo-machines whether they are purely isothermal in operation, purely adiabatic in operation, or a combination of adiabatic/isothermal operation as in turbo-machines employing intercoolers or intercombustion chambers between stages to adjust the temperature of the gas to a given selected temperature prior to introduction of the gas to the next successive stage. And, it has been discovered that the sizing formula of the present invention is also equally applicable to either axial flow or radial flow turbo-machines, and to a wide range of types of turbo-machines including compressors, gas turbines and gas expanders.
Gas expanders are quite similar in construction to gas turbines, but each has a somewhat different emphasis and purpose. In both gas turbines and gas expanders the gas expands as it moves through the several successive stages. However, gas turbines generally have the purpose of generating drive shaft power, for example to power an electrical generator, whereas gas expanders have the principal function of permitting a controlled expansion of gases for the purpose of cooling the gas. Because of the similarity of construction of gas turbines and expanders, the term "gas turbine" as employed hereinafter will include both gas turbines as well as gas expanders, unless otherwise stated.
In one principal aspect of the present invention, a multistage gas turbo-machine includes a first stage and a second stage of differing sizes. Each stage has turbine blades which are contacted by the gas, an inlet in each stage for introducing the gas to the turbine blades in the stage, a discharge from each stage for discharging the gas from the turbine blades in the stage, and the discharge from said first stage communicates with the inlet of the second stage. The first and second stages are substantially identical to each other in design and geometric shape, but the linear dimensions of the second stage differ from those of the first stage substantially in accordance with the formula EQU L=D
where L is the ratio of the linear dimensions of the second stage to the first stage and D is the gas density ratio of the first stage, and ##EQU1##
where P.sub.I is the absolute pressure of the gas entering the first stage, P.sub.O is the absolute pressure of the gas as discharged from the first stage, T.sub.I is the absolute temperature of the gas entering the first stage, and T.sub.O is the absolute temperature of the gas as discharged from the first stage.
In another principal aspect of the present invention, the gas turbo-machine includes a power transmission shaft, and at least some of the turbine blades are coupled to the shaft to rotate with the shaft, and the shaft and the rotating turbine blades of the first and second stages rotate at the same speed.
In still another principal aspect of the present invention, the gas turbo-machine is either an axial flow or a radial flow gas turbo-machine.
In still another principal aspect of the present invention, the gas turbo-machine is a compressor, and the linear dimensions of the second stage are smaller than the linear dimensions of the first stage substantially in accordance with the formula.
In still another principal aspect of the present invention, the first and second stages of the compressor are substantially isothermal.
In still another principal aspect of the present invention, the first stage of the compressor also includes stator blades, and the stator blades include an inlet and outlet for passing a coolant through the blades to cool the gas to the substantially isothermal temperature before the gas is discharged from the first stage.
In still another principal aspect of the present invention, at least the first stage of the compressor is substantially adiabatic.
In still another principal aspect of the present invention, the compressor includes an intercooler between the first stage and the second stage to cool the gas discharged from the first stage before the gas enters the inlet of the second stage.
In still another principal aspect of the present invention, the intercooler cools the gas to substantially the same temperature as the gas introduced to the inlet of the first stage.
In still another principal aspect of the present invention, the gas turbo-machine is a gas turbine, and the linear dimensions of the second stage are larger than the linear dimensions of the first stage substantially in accordance with the formula.
In still another principal aspect of the present invention, the first and second stages of the gas turbine are substantially isothermal.
In still another principal aspect of the present invention, the first stage of the gas turbine also includes a fuel injector which injects fuel into the first stage to heat the gas to the substantially isothermal temperature before it is discharged from the first stage.
In still another principal aspect of the present invention, at least the first stage of the gas turbine is substantially adiabatic.
In still another principal aspect of the present invention, the gas turbine includes a combustor between the first stage and the second stage which heats the gas discharged from the first stage before the gas enters the inlet of the second stage.
In still another principal aspect of the present invention, the combustor heats the gas to substantially the same temperature as the gas introduced to the inlet of the first stage.
In still another principal aspect of the present invention, the gas turbo-machine includes a generator for generating electrical power, and the aforementioned power transmission shaft mechanically couples the turbine blades with the generator.
In still another principal aspect of the present invention, the gas turbo-machine with the generator includes a compressor and a gas turbine, one or both of which includes the aforementioned first and second stages. The gas from the compressor is discharged to the gas turbine, and a heat exchanger (regenerator) is positioned between the compressor and the gas turbine. The discharge from the gas turbine is used to heat the gas being discharged from the compressor before it is introduced to the gas turbine with the heat content of the gas which is discharged from the gas turbine.
In still another principal aspect of the present invention, water is introduced to the first stage of the compressor from below the thermocline of a large body of water, the first stage also includes stator blades, and the stator blades include an inlet and outlet for passing the water through the blades to cool the gas to the substantially isothermal temperature before the gas is discharged from the first stage.
In still another principal aspect of the present invention, a method of designing and constructing a multistage gas turbo-machine comprises preselecting the operating conditions for the gas turbo-machine of gas pressure ratio, gas intake temperature and gas flow rate. A master stage is constructed to have a given design and geometric shape which results in substantially the optimum efficiency during operation of the master stage under the preselected operating conditions. At least one additional stage is then constructed which is substantially identical to the master stage in geometric shape and design, but in which the linear dimensions of the additional stage differ from those of the master stage substantially in accordance with the aforementioned formula.
These and other objects, features and advantages of the present invention will be more clearly understood through a consideration of the following detailed description.